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Slide 3
Earth-Mars-Earth Transportation Pathways
Earth Surface
Earth Orbit
Earth-Mars Transfer Mars-Earth Transfer
Mars Orbit
Mars Surface
Likely primary mission pathway
Unlikely pathway Potential abort pathway
Slide 4
Direct Return
The crew uses the TSH for Earth-Mars transfer and Mars surface stay
Return from Mars surface to Earth in ERV Only feasible if ISPP is utilized because of
large ERV habitat mass Example: Zubrin’s “Mars Direct”
2 vehicle designs required: TSH, ERV Mars surface rendezvous, all assets on the
surface and accessible to crew Split mission with pre-deployment is
required because of the use of ISPP: ERV needs to be fully fueled and ready for
Earth return before crew leaves Earth
Marssurface
Marsorbit
ERV TSH
Required crew transferbetween vehicles
Broken lines: uncrewed operations
Solid lines: crewed operations
Different colors indicatedifferent vehicles
Slide 5
Direct Return + Surface Habitat
Earth-Mars transfer in Transfer Habitat 1
Surface stay in surface habitat (SH)
Return from Mars surface to Earth in transfer habitat (TH-2) deployed one opportunity before
This strategy is only feasible with ISPP because of large return habitat mass
Mission mode requires either a single landing site or a chain of landing sites
2 vehicle designs required: SH and TH
Mars surface rendezvous, all assets on the surface and accessible to crew
Split mission required due to need for pre-deployed TH-2
Makes most sense for a base: surface infrastructure (habitat, mobility, power system) could be emplaced once and re-supplied
Marssurface
Marsorbit
TH-2TH-1SH
Slide 6
Transfer & Surface Hab and MOR
Earth-Mars transfer and Mars surface stay in TSH
Ascent to Mars orbit in MAV Return from Mars orbit to Earth in ERV This strategy could work both with and
without ISPP because the MAV crew compartment could be made very light (>5 mt)
Example: Mars semi-direct, Mars DRM 1.0, 3.0
The Mars staging orbits would likely be dependent on the utilization of ISPP:
ISPP: highly elliptic Mars staging orbit No ISPP: circular low Mars staging orbit
3 vehicle designs required: TSH, MAV, ERV
Mars surface and Mars orbit rendezvous
Split mission with pre-deployment is required in case of ISPP for MAV ascent propellants
Marssurface
Marsorbit
ERV MAV TSH
Slide 7
Mars Descent / Ascent Vehicle and MOR
Earth-Mars and Mars-Earth transfers in Interplanetary Transfer Hab (ITH)
Mars descent and ascent in Mars Descent & Ascent Vehicle (MDAV)
Option for propulsive abort to orbit exists during Mars powered descent
Surface stay in surface hab (SH) This strategy works both with ISPP and
without
The Mars staging orbits would likely be different for ISPP and no ISPP:
ISPP: highly elliptic Mars staging orbit No ISPP: circular low Mars staging orbit
3 vehicle designs required: ITH, MDAV, SH Mars surface and Mars orbit rendezvous Makes most sense for a base: surface
infrastructure (habitat, mobility, power system) could be emplaced once and re-supplied
Marssurface
Marsorbit
ITHSH MDAV
Slide 8
Pre-Deployed Mars Descent / Ascent Vehicle and MOR
Earth-Mars and Mars-Earth transfers in Interplanetary Transfer Hab (ITH)
Mars descent in MDAV-2, Mars ascent in pre-deployed MDAV-1
Surface stay in pre-deployed surface hab (SH)
This strategy would only be chosen if ISPP is utilized for the MDAV
Mission mode requires either a single landing site or a chain of landing sites
The Mars staging orbits would likely be different for ISPP and no ISPP:
ISPP: highly elliptic Mars staging orbit No ISPP: circular low Mars staging orbit
3 vehicle designs required: ITH, MDAV, SH Mars surface and Mars orbit rendezvous Split mission required due to pre-deployed
MDAV Makes most sense for a base: surface
infrastructure (habitat, mobility, power system) could be emplaced once and re-supplied
Marssurface
Marsorbit
ITHSH MDAV-1MDAV-2
Slide 9
Landing & Surface Habitat and MOR
Earth-Mars and Mars-Earth transfers in Interplanetary Transfer Hab (ITH)
Mars landing & surface stay in Landing & Surface Hab (LSH)
Ascent to Mars orbit in MAV
This strategy works both with ISPP and without
The Mars staging orbits would likely be different for ISPP and no ISPP:
ISPP: highly elliptic Mars staging orbit No ISPP: circular low Mars staging orbit
3 vehicle designs required: ITH, LSH, ERV Mars surface and Mars orbit rendezvous
Marssurface
Marsorbit
ITHMAV LSH
Slide 10
Mapping of Major Mars Elements to Mission Phases
Major Mars Systems and Phases ES
->E
O
EO
EO
->E
MT
EM
T
EM
T->
MO
MO
MO
->M
S
MS
MS
->M
O
MO
MO
->M
ET
ME
T
ME
T->
ES
In-space and Surface Habitation x x x x x ? x ? x x xAscent Crew Compartment ? ?Crew Earth Ascent and Entry Vehicle x ? ? xEarth Launch and Departure Propulsion x xMars Aero-systems x xDescent and Landing Systems xMars Ascent and Departure Propulsion x xSurface Exploration Systems xSurface Power Systems xISRU Systems ?
Slide 11
Major Drivers of Selected Mars Systems
Earth Launch and Departure Systems are driven by the total payload which must be launched towards Mars and the Delta-V associated with the required Earth-Mars transfer trajectory.
In-Space Habitation Systems are driven by crew size and the duration of the Earth-Mars transfer trajectory
Mars Aero-Systems for Mars orbit capture and Mars entry and descent are driven by the payloads they must support and the Mars atmospheric entry velocities of the Earth-Mars transfer trajectories they must withstand
Mars Landing Systems are dependent upon the payload they must deliver to the surface and the state (velocity and altitude) at which they must begin operation
Slide 12
Earth Departure Delta-V Standard (no abort option) Conjunction Trajectories
3.50
3.60
3.70
3.80
3.90
4.00
4.10
4.20
4.30
4.40
4.50
4.60
4.70
4.80
4.90
5.00
120 130 140 150 160 170 180 190 200 210 220 230 240 250 260 270
Earth-Mars Transfer Time [Days]
Ear
th D
epar
ture
Del
ta-V
[km
/s] 2020
2022
2024
2026
2028
2031
2033
2035
2037
Slide 13
Earth Departure Delta-V Conjunction Trajectories with Earth-Mars-Earth Abort Option
3.5
3.6
3.7
3.8
3.9
4.0
4.1
4.2
4.3
4.4
4.5
4.6
4.7
4.8
4.9
5.0
0 0.3 0.6 0.9 1.2 1.5 1.8 2.1 2.4 2.7 3
Abort Delta-V [km/s]
Ear
th D
epar
ture
Del
ta-V
[km
/s]
2020
2022
2024
2026
2028
2031
2033
2035
2037
03-Year
Free Return2-Year
Free Return
Slide 15
Abort Type
No Abort
Prop Abort or Hybrid
FRT Abort
Outbound: TSH
Outbound: ERV
Outbound: TSH or ERV
Outbound Abort? Ascent Abort? Outbound Config # Ares I / COTS # Ares V (Base) # EDS (Base) CEV (COTS)
Yes Yes MAV & ERV 0 3 (2) 3 (2) 1
Yes Yes TSH & ERV 0 3 3 1
Yes YesTSH, Earth
Aerocapture1 3 3 2 (1)
Yes NoTSH, Earth
Aerocapture0 3 3 1
Yes YesTSH &
Droppable CEV0 4 3 2
No Yes TSH 1 3 3 2 (1)
No No TSH 0 3 3 1
Abort Dependencies – MOR Architectures
Slide 16
Ares Launch Vehicle Capability
0
10,000
20,000
30,000
40,000
50,000
60,000
70,000
80,000
90,000
100,000
110,000
120,000
130,000
0 500 1,000 1,500 2,000 2,500 3,000 3,500 4,000 4,500 5,000
LEO Departure Delta-V [m/s]
To
tal
Pay
load
[kg
]
1 EDS, 0 mt added in LEO
1 EDS, 10 mt added in LEO
1 EDS, 20 mt added in LEO
1 EDS, 30 mt added in LEO
1 EDS, 40 mt added in LEO
1 EDS, 50 mt added in LEO
1 EDS, 60 mt added in LEO
1 EDS, 70 mt added in LEO
1 EDS, 80 mt added in LEO
1 EDS, 90 mt added in LEO
1 EDS, 100 mt added in LEO
1 EDS, 110 mt added in LEO
1 EDS, 120 mt added in LEO
1 EDS, Max added in LEO
2 EDS, HEO Rendezvous
2 EDS, Staged departure (max)
Single Launch
"1.5 Launch"1
2
35
4
Slide 17
Selected Mars Launch and Earth Departure Configurations
Ares VCore & SRBs
Ares V EDS
Blue = Payload; Gray = EDS Fairing; Red = Nuclear Thermal Stage
1 Ares V, H2/O2 2 Ares V, H2/O2 1 Ares V, NTR 2 Ares V, NTR
Slide 18
Launch and Earth Departure Performance
Trans-Mars Injection Performance vs. Transit Time (Minimum Across Mission Opportunities)
0
10
20
30
40
50
60
70
80
90
100
110
120
120 130 140 150 160 170 180 190 200 210 220 230 240 250 260 270
Earth Mars Transit Time [Days]
TM
I P
ay
loa
d [
mt]
2 Ares V, NTR
1 Ares V, NTR
2 Ares V, H2/O2
1 Ares V, H2/O2
Slide 19
ESAS Launch Vehicle Mars Capability 1 CaLV 1 CLV, 1 CaLV 2 CaLV, 1 EDS 2 CaLV, 2 EDS
~40 mt TMI ~50 mt TMI ~60 mt TMI ~90 mt TMI
~30 mt MO ~37 mt MO ~45 mt MO ~67 mt MO
~20 mt MS ~25 mt MS ~30 mt MS ~45 mt MS
TMI – Trans-Mars Injection; MO – Mars Orbit; MS – Mars Surface
Slide 20
Georgia Tech CE&R Aeroentry Analysis (Ventry = 4.63 km/s, L/D = 0.5)
15 m diameter aeroshell
Aeroshell Sizing Impact on TMI Mass CE&R aerocapture and aeroentry analysis indicated that aeroshell sizing
(diameter) would have a major impact on maximum mass of Mars systems ESAS CaLV has a fairing diameter of 8.4 meters, although larger fairings for Mars
systems would likely be possible For equal ballistic coefficient, the following entry mass limits likely apply for
entry systems of the specified diameter:
8.4^2 = 71 10^2 = 100 12^2 = 144 15^2 = 255
Diameter [m] 8.4 10 12 15
Entry Mass Range [mt] 25 to 31 35 to 44 51 to 64 80 to 100
10 m 15 m12 m8.4 m
Slide 21
Allowable Entry Mass vs. Aeroshell Diameter
0
20
40
60
80
100
120
8 9 10 11 12 13 14 15 16Aeroshell Diameter [m]
TM
I / E
ntr
y M
ass
[mt]
Mach 3 @ ~9 km Altitude
Mach 3 @ ~12 km Altitude
Aeroshell Sizing Impact on TMI Mass CE&R aerocapture and aeroentry analysis indicated that aeroshell sizing
(diameter) would have a major impact on maximum mass of Mars systems ESAS CaLV has a fairing diameter of 8.4 meters, although larger fairings for Mars
systems would likely be possible For equal ballistic coefficient, the following entry mass limits likely apply for
entry systems of the specified diameter:
8.4^2 = 71 10^2 = 100 12^2 = 144 15^2 = 255
10 m 15 m12 m8.4 m Diameter [m] 8.4 10 12 15
Entry Mass Range [mt] 25 to 31 35 to 44 51 to 64 80 to 100
1 LV, 1 EDS
1.5 LV, 1 EDS
2 LV, 1 EDS
2 LV, 2 EDS
Slide 22
Highly Elliptic Mars Orbits
Use of highly elliptic Mars orbits can decrease trans-Earth injection (TEI) delta-V relative to TEI from low Mars orbit
In this analysis, a ~10 hour period highly elliptic orbit is assumed in elliptic orbit cases, which decreases TEI delta-V by 1,000 m/s (this delta-V is added to Mars ascent requirements)
For the above chart, the pericenter is 3,700 km, equivalent to an altitude of ~300 km
Mars Elliptic Orbits (Delta-V from LMO and Period of Elliptic Orbit)
0
20
40
60
80
100
120
140
0 20,000 40,000 60,000 80,000 100,000 120,000
Apocenter [km]
Delt
a-V
[10 m
/s]
an
d P
eri
od
[h
]
Delta-V
Period
Slide 29
Motivation for Solar Power on Mars
Mars missions studies frequently select nuclear fission reactors for providing surface power based upon technical considerations
e.g., mass, complexity, volume, etc. However, requiring nuclear power places a policy constraint on the
critical path for human Mars missions Need sustained funding to develop and test reactor Need political approval for every Mars mission including a reactor
Such a constraint compounds already existing policy challenges Increased dependence on the stance of political parties/politicians Sensitivity to changes in public view-point with respect to nuclear power
in general Other options such as dynamic isotope power systems also suffer
from approval and funding constraints
If solar power can be used as primary power source on Mars it could increase the political feasibility and sustainability of human Mars missions
Note: These policy constraints would also apply to nuclear thermal and nuclear electric propulsion
Slide 30
Daily Solar Incidence Energy Levels (Tracking Arrays, No Atmosphere)
0
2
4
6
8
10
12
0 100 200 300 400 500 600 700
Date in Sols (Perihelion = 0)
kW-h
(so
lar)
/ m
^2 /
sol
Equator
45-degrees North
45-degrees South
Energy Storage Factor Relative to Equator, Sized by Winter Solstice
0
0.5
1
1.5
2
0 5 10 15 20 25 30 35 40 45 50 55 60 65Latitude [deg]
Fac
tor
[-] Storage Capacity Factor
Is Solar Power Technically Feasible?
Preliminary analysis during our CE&R study indicated that solar power would be feasible as the primary power source for human Mars missions
Reasonably straightforward for missions without ISRU, more challenging for missions with extensive ISRU
We will conduct a further study of the technical feasibility and consequences of relying upon solar power for human Mars missions
Factors to be considered include required day and night-time power/energy levels, landing location (latitude), seasons and distance to Sun, atmospheric opacity and light scattering (including statistical nature of dust storms), array type (tracking, fixed-incline, fixed-horizontal)
Relying on solar power would likely place constraints on where missions could be conducted, available power levels, and may drive towards a single base (if a large investment is required to emplace the solar power system)
Slide 31
Winter Solstice Light and Dark Periods by Latitude
0
2
4
6
8
10
12
14
16
18
20
22
24
26
0 10 20 30 40 50 60 70 80 90
Latitude [deg]
Per
iod
[h
ou
rs]
Dark
Light
Energy Storage and Generation Factors Relative to Equator, Sized by Winter Solstice
0
1
2
3
4
5
6
7
8
9
10
0 5 10 15 20 25 30 35 40 45 50 55 60 65
Latitude [deg]
Fac
tor
[-]
Generating Capacity
Storage Capacity
Illumination Fraction over Martian Year for Selected Latitudes
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
0 30 60 90 120 150 180 210 240 270 300 330 360
Solar Longitude, Relative to Winter Solstice [degrees]
Daily
Illu
min
atio
n Fr
actio
n [-]
0 Degrees
10 Degrees
20 Degrees
30 Degrees
40 Degrees
50 Degrees
60 Degrees
70 Degrees
80 Degrees
90 Degrees
Illumination Fraction over Martian Year for Selected Latitudes
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
0 30 60 90 120 150 180 210 240 270 300 330 360
Solar Longitude, Relative to Winter Solstice [degrees]
Daily
Illu
min
atio
n Fr
actio
n [-]
33 Degrees
36 Degrees
39 Degrees
42 Degrees
45 Degrees
48 Degrees
51 Degrees
54 Degrees
57 Degrees
60 Degrees
Slide 34
What to do on Moon to prepare for Mars
Three main areas in which the Moon can offer aid in preparing for Mars, which could be considered objectives for lunar campaigns:
1. Testing systems, technologies, and procedures for Mars exploration in an environment distinct from Earth.
2. Increasing understanding of partial gravity (possibly coupled with radiation) impacts on crew health and performance.
3. Providing an intermediate milestone for human space exploration efforts.
Slide 35
How to measure those things
Metrics for each of the three areas of objectives for Mars preparation using the Moon are as follows:
1. Degree to which lunar systems are similar to Mars systems and fraction of Mars systems validated during lunar activities
2. Number of crew exposed to particular durations of lunar partial gravity (e.g., # at 1 month, # at 3 months, # at 6 months, etc.), with longer durations preferred
3. Date of initial occurrence of high visibility events (e.g., lunar vicinity flight, human lunar landing, long-duration mission, total surface time of Mars mission)
Slide 36
Mars Exploration Elements Following list of elements required for Mars exploration
Earth Launch and Entry Crew Cabin(s) Heavy Lift Launch Vehicle and Earth Departure Systems Descent Stage Heatshields Long-term Surface Habitat Mars Ascent Vehicle (Cabin and Propulsion) Earth Return Vehicle (Habitat and Propulsion) EVA and Mobility Systems Surface Power Systems
Following list of technologies beneficial for Mars missions In-Situ Propellant Production/In-Situ Consumables Production Mars ISRU Compatible Propulsion (e.g., CH4/O2, C2H4/O2)
Items denoted in blue indicate high potential for Moon-Mars commonality
Slide 37
Moon-Mars Exploration System Commonality
Shuttle Operations
Station Operations
Mars Exploration
System Development
Lunar Transportation System Operations
Lunar Long Duration System
Development
Budget Limit
Lunar Long Duration System Operations
Mars Exploration
System Operations
Lunar Transportation System Development
Shuttle Operations
Station Operations
Mars Exploration
System Development
Lunar Transportation System Operations
Lunar Long Duration System
Development
Budget Limit
Lunar Long Duration System Operations
Mars Exploration
System Operations
Lunar Transportation System Development
Shuttle Operations
Station Operations
Lunar Transportation System Operations
Lunar Long Duration System
Development
Budget Limit
Lunar Transportation System Development
Lunar Long Duration System Ops
Mars Exploration
System Operations
Mars Exploration
System Development
Shuttle Operations
Station Operations
Lunar Transportation System Operations
Lunar Long Duration System
Development
Budget Limit
Lunar Transportation System Development
Lunar Long Duration System Ops
Mars Exploration
System Operations
Mars Exploration
System Development
Shuttle Operations
Station Operations
Mars Unique Element
Dev.
Transportation System Operations
Long Duration System
Development
Budget Limit
Transportation System Development
Long Duration System Operations
Mars Unique Element Operations
Shuttle Operations
Station Operations
Mars Unique Element
Dev.
Transportation System Operations
Long Duration System
Development
Budget Limit
Transportation System Development
Long Duration System Operations
Mars Unique Element Operations
Common Moon-Mars Exploration System,Option to Maintain Lunar Missions
Distinct Moon, Mars Exploration Systems,Lunar Missions Curtailed
Distinct Moon, Mars Exploration Systems,Lunar Operations Maintained
If distinct systems are developed for Moon and Mars, we may:
Significantly delay Mars operations Need to curtail lunar operations to enable
Mars (development, operations), resulting in a Moon-Mars mission gap
Never get to Mars at all, because the renewed major investment is not sustainable
By developing a common Moon-Mars exploration system, we can overcome these obstacles and also:
Directly validate key Mars elements during lunar missions
Gain experience in routine production and system operation, decreasing cost and risk
Avoid workforce disruption during transition from Moon to Mars, and possibly continue lunar operations during Mars missions
Provide direct tie between Moon and Mars exploration in the eyes of the public and Congress
Slide 38
Commonality Strategy – Transportation Development Roadmap
Design Philosophy: Maximize hardware commonality to minimize gap between lunar and Mars missions and overall development and production costs
CEV + IPU (27 m3 ):
Integrated aeroshell
Mars Mission Hardware
LEO / ISS Mission Hardware
Common in-space propulsion stage (LCH4 / LOX):Core propulsion stageXL strap-on tanksXXL strap-on tanks (ERV)
Heavy Lift Launch Vehicle:(“2 stages”, 100 mt to LEO)
Short Lunar Mission Hardware
Habitat core and inflatablepressurized tent forplanetary surfaces:
Long Lunar Mission Hardware
Note: Block upgrades across phases are not depicted
LEO propulsion stage:
CEV launch vehicle:
CEV power pack:
LAT for CEV capsule:
SDLV upper stage (125 mt to LEO),potentially EDS-derived:
Mars landing gear & exoskeleton:
Engine 1 (LCH4 / LOX)Restartable, non-throttleable:
Common Earthdeparture stage(LH2 / LOX)
Engine 2 (LCH4 / LOX)Throttleable:
Lunar landing gear & exoskeleton:
Slide 39
Base Moon-Mars Exploration System Commonality Concept
High-level commonality concept developed during Base Period using selected Moon and Mars architectures
Commonality focused on design reuse of complete elements, with modularity in “Yellow Stage” and habitat design
Develop high-level scheme to identify elements where commonality may be beneficial Can be based upon elements with similar capabilities (or requirements) Need to be careful which requirements are compared
e.g., for a propulsion stage, the combination of delta-v, payload, and thrust characterize the capability (to first order); taken in isolation they do not
Develop commonality concept in further detail Trades must be performed between modularity/platforming or “stretchable” options relative to a single design for many use cases
Note: While commonality shown for a particular pair of architectures, approach is not unique to those chosen
Lunar Transportation Architecture Mars Transportation Architecture
Slide 40
Extensible Destination Vicinity Propulsion System
Core LCH4 / LOX propulsion stage
Derived LCH4 / LOX propulsion stages
Exploded viewIntegrated view
-4 non-throttleable LCH4/LOX engines-4 common-bulkhead tanks (CBH)-primary and secondary structure
Mars TEI Mars ascentLunar / Mars descent
-4 non-throttleable LCH4/LOX engines-4 common-bulkhead tanks-4 spherical extension tanks (2 CH4, 2 O2)-primary and secondary structure-dedicated exoskeleton for Mars TEI
-4 non-throttleable LCH4/LOX engines-8 common-bulkhead tanks-primary and secondary structure-additional primary structure
-4 throttleable LCH4/LOX engines-8 common-bulkhead tanks-primary and secondary structure-dedicated Moon and Mars landing gears and exoskeletons
Slide 41
Common Destination Vicinity Propulsion System
Modular solution for Destination Vicinity Propulsion System Common propulsion stage core employed in all use-cases (sized by Lunar Ascent & TEI) Duplicate set of tanks (relative to core) provides additional propellant for Lunar/Mars Descent
and Mars Ascent Extra-large set of strap-on tanks used for TEI from Mars on Earth Return Vehicle Descent stage structural ring and landing gear specific to destination due to distinct loading
conditions Common ascent engines, common descent engines for Moon [2 engines] and Mars [4 engines]
Moon MarsCrew Transport
Surface Habitat
Mars Ascent Vehicle
Transfer and Surface Habitat
Earth Return Vehicle
Slide 42
Moon-Mars Common System Vehicle Stacks
Post-Earth departure commonality mass overhead relative to customized systems:
Lunar Direct Return (Arch 1) Mars Orbit Rendezvous: Combined Trans. and Surf. Habs (Arch. 969)
Short Mission Long MissionLunar Crew
Transfer System
Lunar Long-Duration
Surface Habitat
Outbound Transfer & Surface Habitat
Earth Return Habitat & Propulsion
Mars Ascent Vehicle & Return CEV
81 100 112
112 106
106
106
106
59 39
36 34
9
21
9 mt
27 AS: 33
9
DS: 33DS: 33
Hab: 49
TEIS: 57
Hab: 25
HS: 34 HS: 34HS: 34
Number launches (HLLV+CEVLS):
Elements combine together to form vehicle stacks for variety of missions
Numbers at left represent wet mass in metric tonnes of elements in LEO
Earth Departure Stages have the same dry mass (11 mt) and maximum wet mass (112 mt)
CEVLV capacity 30 mt Lunar HLLV capacity 100 mt Mars HLLV upgraded to 125 mt
Low commonality overhead due to appropriate use of modularity to support variants
2+1 2+0 3+1 3+0 3+0
1% 2% 4% 3% 2%
IMLEO commonality overhead relative to customized systems: 13% 20% 4% 4% 3%
63% savings in unique element dry mass for common vs. custom system design
For modest mass increase, Mars-back commonality offers significant savings in development and production
Slide 43
Extending ESAS Elements to Mars Missions Based upon the significant capability of ESAS launch vehicles to deliver
payloads to Mars, options to extend the remaining elements were assessed
In the baseline architecture presented: Mars-dedicated aeroentry and propulsion systems are developed for
aerocapture/descent, landing, ascent, and trans-Earth injection The CEV is extended to provide habitation during Earth launch and entry, and during
Earth-Mars and Mars-Earth transit in combination with the LSAM-derived Mars Landing and Ascent Vehicle crew compartment, as part of a dual-launch, dual-heatshield crew transportation system
A large surface habitat capable of supporting up to 6 crew members is positioned to the Martian surface prior to the arrival of the crew in a single launch
Single launch logistics flights pre-position consumables, surface exploration equipment, and power infrastructure for initial mission, resupply consumables and spares for subsequent missions
Total number of CaLV launches required (no CLV launches are needed) is presented for a low launch demand and a high launch demand scenario
“Low demand” could be achieved with 4 crew (in one crew transportation system), methane-oxygen propulsion for maneuvers near Mars, and ISRU for both consumables production and propellant production
“High demand” could be achieved either with 4 crew (in two crew transportation systems), hypergolic propulsion, and no ISRU, or with 6 crew (in two crew transportation systems), methane-oxygen propulsion, and no ISRU
While not presented, intermediate launch demand options also exist with other crew size and technology combinations
Slide 44
Mars Crew Transportation System Concept
Launch Configuration Trans-Mars ConfigurationLogistics Flights Surface Habitat Crew Transport
Mars Orbit
Interplanetary Transfer
Earth Vicinity
Mars Surface
Conceptual Mars Exploration Architecture Based on Lunar Elements
Slide 45
Logistics Flights Surface Habitat Crew Transport
Conceptual Mars Exploration Architecture Based on Lunar Elements
Mars Orbit
Interplanetary Transfer
Earth Vicinity
Mars Surface
# of CaLVs by Mission
Initial Subs.Log. and Infra. 3 1Crew Transport 2 2
Total CaLVs 5 3
Initial Subs.Log. and Infra. 4 2Crew Transport 4 4
Total CaLVs 8 6
Low Demand
High Demand
Low demand: 4 crew, methane-oxygen prop, ISRU
High demand: 4 crew, hypergolic prop, no ISRU OR 6 crew, methane-oxygen prop, no ISRU
Slide 46
Mars Considerations for the Moona.k.a. Mars-Back to the Moon
Additional Considerations for building and extensible moon/Mars exploration system
Slide 47
What to do on Moon to prepare for Mars
Areas in which Moon can aid in preparing for Mars:
1. Testing systems, technologies, and procedures for Mars exploration in an environment distinct from Eartha) Gain operational and developmental experience;
test operations procedures
b) Test technologies for Mars systems (e.g., method of water recycling
c) Test Mars subsystems (e.g., ECLSS, fuel cell)
d) Test Mars systems (e.g., habitat, rover)
2. Increasing understanding of partial gravity (possibly coupled with radiation) impacts on crew health and performance
3. Providing an intermediate milestone for human space exploration efforts
Metrics for each area:1. Degree to which lunar systems,
subsystems, etc. are similar to Mars systems and fraction of Mars systems validated during lunar activities (with validated earlier being better)
2. Number of crew exposed to particular durations of lunar partial gravity (e.g., # at 1 month, # at 3 months, # at 6 months, etc.), with longer durations preferred
3. Date of initial occurrence of high visibility events (e.g., lunar vicinity flight, human lunar landing, long-duration mission, total surface time of Mars mission)
Slide 48
Measuring Readiness for Mars
It may be useful to define a “Mars Exploration Readiness Level” (MERL) to determine Mars preparedness
Would serve a similar function to the TRL scale for measuring technologies
Having such a scale would enable: Reviewing status Monitoring progress Determining criteria for tests
Lunar testing would be one means of increasing MERL Laboratories, terrestrial Mars-analogs, LEO, deep-space,
and Mars could also serve as test environments
MERL could be applied to areas 1 and 2 from prior listing (i.e., technical readiness and human physiological readiness)
MERL could be a hierarchical measure, and at top level would depend upon overall Mars architecture(s) under consideration
ME
RL
Time
Level Required by NASAto commence Mars missions
Slide 49
Potential Areas of Moon-Mars Commonality Habitat sub-systems Habitat structure
Note: Mars mission will require relatively large habitat volume, should limit surface assembly due to safety considerations
Mars EDL system would place additional constraints on habitat dimensions, CG Surface exploration systems (i.e., rovers, EVA aids, etc.) EVA suits
Possible differences in thermal, note also different in-situ resources may impact design Power systems (generation and storage)
For solar, storage is easier on Mars, generation harder Operations commonality (time lag and associated impacts)
Should lunar missions operate in this manner from the start? Resupply and logistics? Propulsion?
Ascent, TEI: LCH4/LOX, C2H4/LOX, RP-1/LOX?, etc. H2/LOX for descent?
Note: Lunar ISRU is intentionally not on this list
While these areas offer potential for commonality, the list does not indicate what should be done in designing lunar systems to make them common with Mars
Slide 50
Mars Considerations for Lunar Systems
In order to determine Mars-related considerations for the development of lunar systems (so as to aid in lunar system design decision making), we ask:
How is Mars different from the Moon?
How do the differences tend to impact Mars exploration systems?
What can be done with lunar systems to have higher relevance towards Mars systems?
Slide 51
Differences between Mars and the Moon (1)
Mars has an atmosphere Mars systems require an aeroshell for EDL
Aeroshell must be launchable at Earth (constrains diameter) Integrated system must be stable through atmospheric flight Tends to lead to compact, low cg descent stage and payloads for Mars For Moon: Consider designing payloads to meet Mars aeroshell constraints
CO2 is easily available on Mars Mars systems would tend to use atmospheric-based ISRU, rather than more
challenging, complex, and risk regolith-based ISRU For Moon: Consider use of Mars ISRU related processes in ECLS and power
systems (e.g., water electrolysis, Sabatier) For Moon: Do not justify lunar regolith or water-ice based ISRU on the basis
of relevance to Mars; do not predicate Mars ISRU on success with lunar ISRU
Mars surface radiation environment more benign than Moon Mars systems may require less shielding while on surface For Moon: Unclear if this carries implications; may impact common mobility
system design in that Mars systems would not need as much shielding
Slide 52
Differences between Mars and the Moon (2)
Campaign- / mission-level differences Mars missions will be longer and possibly feature larger crew
Required surface habitat volume will be greater for Mars than Moon For Moon: Consider oversized (relative to lunar requirements) habitat or
means for increasing habitat volume between Moon and Mars
Full Mars habitation capabilities are required from the start Mars systems would tend to limit assembly required to provide crew
habitation (assembly might be used to augment capabilities) For Moon: Consider means of providing core crew habitation functionality
without assembly
Mars missions do not have any opportunity for re-supply Leads towards pre-positioning of required consumables and sufficient up-
front planning to anticipate full needs of mission For Moon: Consider logistics strategy that does not rely on rapid re-supply,
but uses pre-placement or pre-planned re-supply
Slide 53
Differences between Mars and the Moon (3) Mars has approximately twice the gravity of the Moon
Mars surface operations will be under a higher gravitational load than lunar surface operations Tends towards desire for lighter space suits for Mars Tends towards decreased reliance on astronauts lifting, carrying items Mobility energy requirements will be higher for a given mass Mobility systems may become stuck more easily / more stuck on Mars For Moon: Consider means to decrease space suit mass further than may be needed for
lunar ops; limit degree to which astronauts lifting is required For Moon: Aim to minimize mobility “payload” mass; consider means to recover higher
weight mobility systems
Mars is further away from Earth than the Moon Mars missions will feature a significant round-trip communications lag
Mars missions operations will conducted in very different manner (relative to operations to date) in order to accommodate time lag; crew will have significantly greater autonomy, with Earth supporting and advising on a longer time-scale
Mars systems will require greater automation in order to monitor system health and limit workload on surface crew
For Moon: Consider means of simulating Earth-Mars communications lag in lunar operations; establish lunar surface operations methodology including time lag
Slide 54
Differences between Mars and the Moon (4) Mars is further from Sun and has a ~24 hour day
Solar incidence is lower although eclipse time is shorter Solar power generation more challenging on Mars Energy storage (e.g., regen fuel cells) is easier on Mars For Moon: Consider means to modularize lunar solar power generation system so that
additional units can be included in Mars case; consider means to decrease energy storage capacity for Mars relative to Moon
Note: Nuclear power is not necessarily required for Mars
Mars thermal environment is different from the Moon Regolith temperature more moderate on Mars than Moon, also have
atmospheric interactions on Mars Different thermal systems might be required between Moon and Mars (although more
analysis needed) For Moon: Identify how relevant lunar systems are towards Mars, determine options for
transitioning from lunar thermal environment to Mars
Mars may have or may have had indigenous life Should restrict contamination of sites of scientific interest with
organisms transported by humans; limit exposure of crew to martian organisms Techniques for sterile sampling of biologically interesting sites may be required Techniques for preventing biological transfer into crewed areas may be required For Moon: Consider testing of such techniques as part of lunar operations
Slide 55
Concept to implement delay in Earth-Moon communications so that lunar surface missions are conducted in a manner and with an organization that is relevant to Mars
Delay would be in Earth to Moon comm. leg (not Moon to Earth), so that mission support could actively assess situation and switch to real-time comm. in an emergency
Primary role of mission support organization would be to assist crew and monitor longer time-scale processes – real time decision-making would reside on the lunar surface
Option to include an Earth based “artificial artificial intelligence” organization may be worthwhile to reduce need to develop onboard automation prior to lunar missions
Would be staffed by personnel simulating the role of “automation”
Should be tightly tied to organization developing automation software, to inform the development of software and test various techniques
Human involvement in this area would be decreased / eliminated over time
Could also have 2 “virtual astronauts” on Earth with real-time comm. links
Could assist in difference in crew size between Moon and Mars
Could monitor EVAs, teleoperate rovers, etc. Could test various roles for crew members in a Mars
mission
Mars-Driven Lunar Missions Operations Concept
MissionSupport
4 to 20 min. lag
4 to 20 min. lag
Slide 56
Concept to implement delay in Earth-Moon communications so that lunar surface missions are conducted in a manner and with an organization that is relevant to Mars
Delay would be in Earth to Moon comm. leg (not Moon to Earth), so that mission support could actively assess situation and switch to real-time comm. in an emergency
Primary role of mission support organization would be to assist crew and monitor longer time-scale processes – real time decision-making would reside on the lunar surface
Option to include an Earth based “artificial artificial intelligence” organization may be worthwhile to reduce need to develop onboard automation prior to lunar missions
Would be staffed by personnel simulating the role of “automation”
Should be tightly tied to organization developing automation software, to inform the development of software and test various techniques
Human involvement in this area would be decreased / eliminated over time
Could also have 2 “virtual astronauts” on Earth with real-time comm. links
Could assist in difference in crew size between Moon and Mars
Could monitor EVAs, teleoperate rovers, etc. Could test various roles for crew members in a Mars
mission
Mars-Driven Lunar Missions Operations Concept
MissionSupport
Artificial Automation
Virtual Astronauts
8 to 40 min. lag
emer
gency
com
m.